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Abstract:

An engine has a blade stage and a circumferential array of blade outer
air seal segments. A support ring carries the blade outer air seal
segments. The support ring has a low-CTE member and a high-CTE member
intervening between the blade outer air seal segments and the low-CTE
member.

Claims:

1. An engine comprising: a blade stage; a circumferential array of blade
outer air seal segments; and a support ring carrying the blade outer air
seal segments and comprising: a non-metallic member; and a metallic
member intervening between the blade outer air seal segments and the
non-metallic member.

2. The engine of claim 1 wherein the metallic member and non-metallic
member have dimensions and physical properties so that: over an
operational temperature range there are distinct stages in which the
circumferential thermal expansion of one member versus the other dictate
radial expansion of the circumferential array of blade outer air seal
segments.

3. The engine of claim 2 wherein the distinct stages include: a stage in
which the circumferential thermal expansion of the non-metallic member
alone essentially dictates said radial expansion; and a stage in which
the circumferential thermal expansion of the metallic member as resisted
by the non-metallic member essentially dictates said radial expansion.

4. The engine of claim 3 wherein the distinct stages include: a stage in
which the circumferential thermal expansion of the metallic member alone
essentially dictates said radial expansion.

5. The engine of claim 1 wherein: the blade stage is a turbine stage.

6. The engine of claim 1 wherein: the metallic member comprises a
nickel-based superalloy; and the non-metallic member comprises a ceramic
matrix composite.

7. The engine of claim 1 wherein: the metallic member comprises an
integral full hoop; and the non-metallic member comprises an integral
full hoop.

8. The engine of claim 1 wherein: the metallic member has a CTE of
3.0-4.0 ppm/C; and the non-metallic member has a CTE of 0.5-2.0 ppm/C.

9. The engine of claim 1 wherein: the metallic member comprises a
circumferential array of edge-to-edge segments; and the non-metallic
member comprises a full hoop.

10. The engine of claim 9 wherein each of the edge-to-edge segments
comprises: a forward half; and an aft half secured to the forward half to
capture the non-metallic member and capture an associated one of the
blade outer air seal segments.

11. The engine of claim 1 wherein the ring further comprises: a radially
compliant member between the metallic member and the non-metallic member.

12. The engine of claim 11, wherein: the radially compliant member
comprises a plurality of springs.

13. The engine of claim 1 wherein the metallic member and non-metallic
member have dimensions and physical properties so that: as a temperature
increases from a first temperature of less than 200 C to a second
temperature of at least 550 C there are distinct stages in which the
physical properties of one member versus the other dictate radial
expansion of the circumferential array of blade outer air seal segments.

14. A method for operating the engine of claim 1 comprising: running the
engine in a range from a first condition to a second condition, a
characteristic inner diameter of the blade outer air seal array
increasing from a first diameter to a second diameter due to thermal
expansion of the support ring wherein: in a first portion of the range,
the expansion is dictated essentially by one of the metallic member and
non-metallic member but not the other; and in a second portion of the
range, the thermal expansion is dictated substantially by the other.

15. The method of claim 14 wherein: the first portion of the range is at
lower temperature than the second portion of the range; in the first
portion of the range, the metallic member essentially dictates the
thermal expansion and a radial gap between the metallic member and
non-metallic member decreases with increasing temperature; and in the
second portion of the range, the thermal expansion is substantially
dictated by the non-metallic member.

16. The method of claim 14 wherein: in the first portion of the range, an
edge-to-edge clearance between segments of the metallic member decreases
with temperature and the non-metallic member essentially dictates the
thermal expansion; and in the second portion of the range, the
edge-to-edge clearance has closed.

17. A method for operating the engine of claim 1 comprising: an
acceleration of the engine from a first speed to a second speed during
which segments of the metallic member engage each other; and a
deceleration of the engine from the second speed to the first speed
during which the segments of the metallic member decouple from each other
at a larger radius that a radius at which they originally engaged each
other; and a reacceleration wherein the segments of the metallic member
reengage each other while still at a larger radius that the radius at
which they originally engaged each other.

18. An engine comprising: a blade stage; a circumferential array of blade
outer air seal segments; and a support ring carrying the blade outer air
seal segments and comprising: a full annulus member; and a segmented
member intervening between the blade outer air seal segments and the full
annulus member.

19. The engine of claim 18 wherein: in a first portion of an operational
range, an edge-to-edge clearance between segments of the segmented member
decreases with temperature and the full annulus member essentially
dictates the thermal expansion; and in a second portion of the range, the
edge-to-edge clearance has closed.

20. The engine of claim 18 wherein: the segmented member comprises a
nickel-based superalloy; and the full annulus member comprises a CMC or a
gamma titanium.

Description:

BACKGROUND

[0001] The disclosure relates to blade clearance in turbomachinery. More
particularly, the disclosure relates to control via thermal properties of
shroud support rings.

[0002] Gas turbine engines may contain rotating blade stages in fan,
compressor, and/or turbine sections of the engine. Clearance between
blade tips and the adjacent non-rotating structure may influence engine
performance. Clearance may be influenced by mechanical loading (e.g.,
radial expansion of the blades and/or their supporting disks due to
speed-dependent centrifugal loading) and thermal expansion (e.g., of the
blades/disks on the one hand and the non-rotating structure on the
other).

[0003] The high temperatures of the turbine section(s) make clearance
issues particularly significant due both to: (1) the greater significance
of thermal expansion; and (2) temperature-induced modulus reduction which
exacerbates expansion from mechanical loading. In multi-spool engines,
this will be particularly significant in the high speed/pressure turbine
section of the engine. This may be particularly significant in the
engines of combat aircraft which may be subject to greater and more rapid
variations in speed and other operating conditions than are the engines
of civil aircraft.

[0004] Accordingly, a variety of clearance control systems have been
proposed.

[0005] To provide active control, many proposed systems form the
non-rotating structure with a circumferential array of blade outer air
seal (BOAS) segments mounted for controlled radial movement (e.g., via
actuators such as electric motors or pneumatic actuators). An aircraft or
engine control system may control the movement to maintain a desired
clearance between the inner diameter (ID) faces of the BOAS segments and
the blade tips.

[0006] Additionally, various proposed systems have involved tailoring the
physical geometry and material properties of the BOAS support structure
to tailor the thermal expansion of the support structure to provide a
desired clearance when conditions change. Such thermal systems may be
passive. Alternatively, such thermal systems may involve an element of
active control such as via controlled direction of cooling air to the
support structure.

[0007] Proposals for thermal expansion-based systems have included systems
wherein the BOAS support structure comprises two distinct materials
having different coefficients of thermal expansion (CTE) and dimensioned
and positioned relative to each other to provide a staged expansion
wherein the relative influence of each of the two materials changes over
the range of operation. One example of such a system is found in U.S.
Pat. No. 5,092,737.

SUMMARY

[0008] One aspect of the disclosure involves an engine having a blade
stage and a circumferential array of blade outer air seal segments. A
support ring carries the blade outer air seal segments. The support ring
has a low-CTE (e.g., nonmetallic member) and a high-CTE (e.g., metallic
member) intervening between the blade outer air seal segments and the
low-CTE member.

[0009] In various implementations, the metallic member and non-metallic
member have dimensions and physical properties so that over an
operational temperature range there are distinct stages in which the
circumferential thermal expansion of one member versus the other dictate
radial expansion of the circumferential array of blade outer air seal
segments.

[0010] The details of one or more embodiments are set forth in the
accompanying drawings and the description below. Other features, objects,
and advantages will be apparent from the description and drawings, and
from the claims.

[0018] Like reference numbers and designations in the various drawings
indicate like elements.

DETAILED DESCRIPTION

[0019] FIG. 1 schematically illustrates an exemplary gas turbine engine 10
including (in serial flow communication from upstream to downstream and
fore to aft) a fan section 14, a low-pressure compressor (LPC) section
18, a high-pressure compressor (HPC) section 22, a combustor 26, a
high-pressure turbine (HPT) section 30, and a low-pressure turbine (LPT)
section 34. The gas turbine engine 10 is circumferentially disposed about
an engine central longitudinal axis or centerline 500. During operation,
air is: drawn into the gas turbine engine 10 by the fan section 14;
pressurized by the compressors 18 and 22; and mixed with fuel and burned
in the combustor 26. The turbines 30 and 34 then extract energy from the
hot combustion gases flowing from the combustor 26.

[0020] In a two-spool (two-rotor) design, the blades of the HPC and HPT
and their associated disks, shaft, and the like form at least part of the
high speed spool/rotor and those of the LPC and LPT form at least part of
the low speed spool/rotor. The fan blades may be formed on the low speed
spool/rotor or may be connected thereto via a transmission. The
high-pressure turbine 30 utilizes the extracted energy from the hot
combustion gases to power the high-pressure compressor 22 through a high
speed shaft 38. The low-pressure turbine 34 utilizes the extracted energy
from the hot combustion gases to power the low-pressure compressor 18 and
the fan section 14 through a low speed shaft 42. The teachings of this
disclosure are not limited to the two-spool architecture. Each of the
LPC, HPC, HPT, and HPC comprises interspersed stages of blades and vanes.
The blades rotate about the centerline with the associated shaft while
the vanes remain stationary about the centerline.

[0021] FIG. 2 shows one of the stages 60 of blades 62 and its associated
array 64 of blade outer air seal (BOAS) segments 66 (e.g., of the HPT).
The segments locally bound the radially outboard extreme of the core
flowpath through the engine. The segments are supported on a support ring
structure 70. The support ring is subject to thermal expansion. With a
monolithic support ring, the coefficient of thermal expansion (CTE) of
the ring material will influence radial thermal expansion of the ring,
and thus of the inner diameter (ID) faces of the BOAS segments. Such
expansion, relative to combined thermal and centrifugal expansion of the
associated rotor at the blade tips, dictates the change in tip clearance.
High CTE materials would include alloys such as nickel-based superalloys
(e.g., Inco 718). These have an exemplary CTE range of at least 3.0 ppm/C
(more narrowly 3.0-4.0 ppm/C or 3.5-4.0 ppm/C). Low CTE materials include
ceramics and ceramic matrix composites (CMC) and have an exemplary CTE
range of up to 2.5 ppm/C (more narrowly, up to 2.0 ppm/C or 0.5-2.0 ppm/C
or, for exemplary CMCs, 0.5-1.0 ppm/C). As a practical matter, the low
CTE material will have a much lower thermal conductivity than the high
CTE material. Thermal conductivity will have a substantial influence in
view of transient behaviors. For example, with rapid changes in engine
speed/power, centrifugal effects on tip radius may occur before a
corresponding temperature change has full influence on the ring. The
delay is more significant with lower conductivity rings. For example, a
rapid acceleration or deceleration may cause departures from the
equilibrium situation. There are other metallic low-CTE materials that
could be used. One example is gamma titanium. An exemplary CTE for this
is about 2 ppm/C.

[0022] For example, combat aircraft may be subject to rapid acceleration
from cruise conditions. Evidencing the complexity of the problem, such an
acceleration could be from a steady-state cruise condition or could be a
reburst wherein the engine had been operating close to full speed/power
long enough for temperature to depart from equilibrium cruise conditions
whereafter the engine decelerates back to a cruise speed and before the
engine can re-equilibrate, reaccelerates. Accordingly, the engine may be
designed with anticipated non-equilibrium situations in mind.

[0023]FIG. 6 shows simplified plots of tip clearance versus engine power
at fully equilibrated or steady-state conditions. A first plot 550
reflects a hypothetical high CTE ring. A second plot 552 reflects a
hypothetical low CTE ring. An exemplary lowest illustrated power level
PC may be an idle level. An exemplary highest illustrated power
level PH may be a take-off power level. FIG. 6 further shows an
exemplary cruise power level PC. The particular plots 550 and 552
illustrated for the two ring properties are dictated by transient
behaviors. In general, the tip clearance will decrease with power
because: (1) the rotor is subject to both thermal expansion and
centrifugal/inertial expansion, whereas the ring is subject only to
thermal expansion; and (2) the exemplary blade and disk materials have
relatively high CTE so there is not much chance to use a sufficiently
higher CTE ring material to counter the centrifugal/inertial expansion.

[0024] If no transients were involved, the rings could be sized so that
there was essentially zero clearance at the maximum anticipated power. To
the extent that the high CTE ring would tend to thermally expand at a
rate closer to the thermal expansion rate of the rotor, it would have
much lower/tighter clearance as power decreased from maximum compared
with a low CTE ring. However, transient behavior imposes different
requirements on low CTE rings versus high CTE rings. One example of a
transient situation is a rapid deceleration from take-off power followed
by a reacceleration. During the rapid deceleration, the low CTE ring will
contract more slowly than the high CTE ring (due to the associated lower
thermal conductivity). There must be sufficient steady-state/equilibrium
clearance at high power to compensate for the contraction of the ring
during this important transient period. Thus, the steady-state clearance
of the high CTE ring must be greater at high power conditions than that
of the low CTE ring. Thus, if the high CTE plot 550 were selected to be
lower (e.g., of similar slope but having the same high power steady-state
clearance as the low CTE plot) then there would be a pinch or rub
situation in the transient. Assuming abradable coatings or materials,
such a pinch/rub would promptly abrade material to increase clearance and
reset the plot 550 to the higher level illustrated.

[0025] These two exemplary clearances are shown in FIG. 6 as TCHH and
TCLH, respectively. Working back to low power from this condition,
the thermal conductivities of the rings dictate the slopes of the plots
550 and 552. The resulting situation is that the plots 550 and 552
intersect at an intermediate power (in the exemplary case near cruise but
not necessarily so). At steady-state/equilibrium conditions, there are
similar clearances and thus similar efficiencies at this intermediate
power. At higher power, the clearance associated with the high CTE ring
is greater and thus there is less efficiency than with the low CTE ring.
Similarly, at lower power, the clearance is greater with the low CTE ring
and thus there is lower efficiency with the low CTE ring than with the
high CTE ring.

[0026] However, as is discussed below, there may be several opportunities
for using a hybrid ring which includes both high CTE material and low CTE
material. During different stages of operation, expansion of the hybrid
ring is influenced in substantially differing proportions from the two
(or more) materials. As one example, plot 554 shows a hypothetical system
which has three stages of operation: a low power stage 554-1 up to a
power P1 has behavior relatively heavily influenced by the low CTE
material; an intermediate power range 554-2 has behavior more heavily
influenced by the high CTE material; and an upper power range 554-3 above
a power P2 has behavior influenced by both (e.g., reflecting a
weighted average CTE of the two CTE materials; additionally the relative
hoop stiffness of the low CTE ring 72 in tension, and compressive
stiffness of the high CTE carriers 92, will further determine the
behavior of the system. In the illustrated example the slope is about
average of the slope associated with the low CTE material and the slope
associated with the high CTE material.

[0027] For example, FIG. 7 shows a more detailed transient operational
condition. The plot 558 shows the tip radius versus time. At an initial
time t0 the tip radius is at its idle condition RTI. At
t1, the engines are throttled up to take-off power. In a very brief
interval later, at t2, the rotor is at full speed but has not
reached its peak temperature. Accordingly, based largely upon the
inertial/centrifugal forces the radius has expanded to RT2. As the
engine heats, the rate of thermal growth continues gradually decreasing
as the engine approaches a steady-state take-off power condition. In this
example, at t3, as the engine nears what would be a steady-state
take-off power condition, the tip radius is RT3. Power is then
rapidly dropped to idle power. There is initially a very sharp drop in
radius to RT4 at an approximate time t4 wherein much of the
initial/centrifugal component is eliminated (although not all because at
higher temperature the effect of the centrifugal component is greater
than it was at lower temperature). As the rotor cools, there is a further
decrease in radius until RTI is again reached at approximately
t5.

[0028] In a variant transient, however, at time t6 shortly after
t4, there is a rapid reacceleration (plot 558') which occurs almost
instantaneously (e.g., its beginning and end times are not separately
marked). This expands the tip radius up to RT5. Thereafter, the
remaining thermal expansion will bring the tips to the steady-state
take-off power radius at or slightly beyond RT3.

[0029] FIG. 7 also shows plots 560, 562, and 564 for the high CTE ring,
low CTE ring, and hybrid ring, respectively responsive to the power
profile plot 558. Due to the relatively quick cooldown after t5, at
t6, the ring radius of the high CTE ring has shrunk sufficiently to
have pinch with the reacceleration of plot 558'. The hybrid ring can
avoid such pinch as is discussed below.

[0030] A first example of a hybrid support ring 70 comprises two sub-units
or members. One member 72 comprises or consists essentially of a low-CTE
non-metallic member which forms an integral full hoop. More particularly,
the exemplary member 72 forms a continuous (e.g. continuous
microstructure rather than segments mechanically attached to each other)
full hoop. The exemplary member 72, or at least its full hoop portion, is
formed as a ceramic matrix composite. The exemplary member 72 has a
generally rectangular axial/radial cross-section (e.g., a rectangle with
rounded corners) formed by an inboard (radially) or inner diameter (ID)
surface or face 74, an outboard (radially) or outer diameter (OD) surface
or face 76, a forward/fore/upstream surface or face 78, and a
rear/aft/downstream surface or face 80 (FIG. 3).

[0031] The exemplary CMC of the member 72 comprises a collection of
silicon carbide fibers and mono-filiment carbon/silicon carbide fibers
preferentially woven in the hoop direction, with silicon carbide fibers
woven in the axial and radial directions to create a fibrous pre-form. An
interface coating on the fibers, (e.g., primarily boron nitride), may be
applied to impart a weak interface bond. A glass-based matrix may be
injected or hot-pressed into the fibrous perform to create a consolidated
ring of essentially rectangular cross section. An external coating, such
as an environmental barrier coating, may be applied to the exposed
surfaces.

[0032] Another sub-unit or member of the ring 70 is formed by a
circumferential array 90 (carrier ring) (FIG. 2) of BOAS carriers 92
(FIG. 5). In the exemplary configuration, each BOAS carrier 92 carries a
single associated BOAS segment 66. This is discussed further below. The
carriers 92 mount the BOAS segments to the member 72. The carriers are
formed of a high-CTE material (e.g., metal such as a nickel-based
superalloy, although not necessarily at the high end of CTE for such
alloys).

[0033] Each BOAS segment 66 comprises a main body 100 (FIG. 4) having an
ID face 102 and an OD face 104. The main body 100 extends from a
fore/forward/upstream end 106 to a rear/downstream/aft end 108 and has a
first circumferential end/edge surface/face 110 (FIG. 2) and a second
circumferential end/edge surface/face 112. Each of the circumferential
ends 110 and 112 may bear a seal slot 114. With the array assembled,
adjacent circumferential ends 110 and 112 of adjacent segments 66 come
into facing alignment with each other and may receive corresponding edge
portions of a seal 120 (e.g., a metallic feather seal). Outboard of the
OD face 104, the segment 66 may include fore and aft mounting features
130 and 132 (FIG. 5). Exemplary mounting features are lugs or ears having
a radially outward projecting proximal or leg portion 134 and an axially
projecting distal portion 136 (e.g., projecting axially outward (forward
for the forward lug 130 and aft for the aft lug 132)). Each exemplary
segment may have a pair of such fore and aft lugs along each of the
circumferential ends.

[0034] An exemplary BOAS segment may be formed of a cast nickel-based
superalloy. The segment may have an internal cooling passage system (not
shown) and may have a thermal barrier coating (not shown) (at least along
the ID face). The exemplary BOAS segment may represent any of a number of
known or yet-developed BOAS segment configurations.

[0035] As is discussed further below, the feather seal 120 (FIG. 2) spans
a circumferential gap 122 (having a central radial/axial plane 502)
between adjacent BOAS segments. The exemplary gap 122 has a width W1
(e.g., measured as the minimum circumferential width at a given
condition).

[0036] Each exemplary carrier 92 comprises a body (FIG. 4) formed as the
assembly of a fore/front half 140 and an aft/rear half 142. The fore and
aft halves may be secured to each other by one or more fasteners 144. The
exemplary fastener 144 is formed as a single socket-head bolt.

[0037] In the exemplary carrier 92, the fore and aft halves 140 and 142
are generally symmetric across a transverse mating plane 504 which may
form a transverse centerplane of the member 72. For example, the halves
may at least depart from mirror images of each other by the presence of
differing: fastening features (e.g., for cooperation with the bolt 144);
and/or features for registering the halves with each other (e.g., lugs
and mating pockets). Exemplary halves have respective faces 150 and 152
(FIG. 3) along the plane 504 abutting each other. The assembled carrier
halves generally surround/encircle the axial cross-section of the member
72 to grasp the member 72 and secure the carrier to the member 72. Each
exemplary carrier half comprises a portion 160 generally immediately
radially inboard of the member 72 and wrapping slightly around to the
adjacent fore or aft face of the member 72. Each portion 160 has an
associated first circumferential end 162 (FIG. 2) and second
circumferential end 164. Exemplary ends 162 and 164 form circumferential
ends/extremes of the carrier so that the end 162 of each carrier is
separated from the end 164 of the next adjacent carrier by a gap 170
having a width W2.

[0038] For capturing the BOAS segment, the exemplary carrier halves have
respective fingers 180 and 182 (FIG. 4). Each half has a pair of such
fingers 180 or 182 spaced slightly circumferentially inward of the ends
162 or 164. Each finger has a proximal portion 190 extending
longitudinally outward from a root portion adjacent thereto. An
intermediate portion 192 extends radially inward from the proximal
portion 190. A distal portion 196 extends axially inward from the inboard
end of the intermediate portion. Each of these fingers thus interfits and
cooperates with an associated mounting lug 130 or 132 of the blade outer
air seal.

[0039] A central inboard portion of each half has fastening features
cooperating with the fastener 144 for securing the two halves together.
For example, a central inboard portion 200 (FIGS. 3 and 4) of the first
half 140 has a counterbored compartment 201 receiving the head 202 and an
intermediate portion 203 of the fastener 144 while the corresponding
portion 204 of the second half has a threaded bore 206 receiving a
threaded end portion 207 of the fastener. The exemplary BOAS segment
bridges these central inboard portions 200, 204 with mounting lugs 130 or
132 falling on either circumferential side of the associated central
portion. This bridging circumferentially retains the BOAS segment to the
carrier.

[0040] In operation, the engine heats up. As the engine heats up, its
components thermally expand due to their coefficients of thermal
expansion (CTE). FIG. 2 shows tips 220 of the blades 62 at a tip radius
RT from the engine centerline 500. The ID faces 102 of the BOAS
segment bodies are at a radius RS so that the difference between
RS and RT is a tip clearance shown as TC. FIG. 3, further,
shows a radial gap 230 between the ID face 74 of the ring 72 at a radius
RR from the centerline and the adjacent inner face 238 of the
aperture in the carrier receiving the ring. This gap has a radial span of
GR1. The separateness of the carrier segments will otherwise allow
the gap 230 to close when the associated segments shift radially outward.
To bias the gap open, an exemplary compliant member 240 (FIG. 4) is shown
as a circumferential segment of metallic wave spring received in a
radially-inwardly extending compartment in the carrier and engaging the
ring surface 74. As is discussed below, sufficient force may, however,
compress the spring 240 and close the gap 230.

[0041] The exemplary array 64 of BOAS segment 66 along with the support
ring structure is supported/carried by the engine case 250 (FIG. 4). To
do this, each carrier has a pair of lugs shown as a front/fore/forward
lug 252 (FIG. 2) and a rear/aft lug 254 (FIG. 4). The respective fore and
aft lugs extend from the respective fore and aft halves of the carrier
body. The lugs are received in respective channels 256, 258 in
radially-inwardly extending webs 260, 262 of the case 250 extending
radially inward from a sidewall 264 of the case. FIG. 2 shows lateral
circumferential gaps 270 between sides 272 and 274 of each lug 252 and
associated sides 276, 278 of the channel. These are exaggerated. They may
be just small enough to avoid binding over the operational range while
maintaining concentricity of the support ring and the engine case as well
as rotor axis/centerline 500. Additionally, a radial gap 280 is shown
between an outboard/OD end 282 of the lug and a base 284 of the channel.
The gap 280 has a radial span GR2 (FIG. 4).

[0042] The engine may have a characteristic temperature which will
generally increase with engine speed and power. An exemplary temperature
range may be characterized from a low of TO (e.g., at or below 200
C) to a high of TH (e.g., at least 550 C). As the temperature of the
member 72 increases, it will circumferentially expand, thereby, causing a
corresponding radial expansion. This radial expansion will tend to at
least partially counter any thermal and centrifugal expansion of the
blade/rotor system which would close the tip clearance gap. An initial
stage of such expansion is shown by 554-1 in FIG. 6. Due to the
relatively low thermal expansion of the CMC material of the member 72,
its expansion may be insufficient to fully counter blade/disk expansion.
Accordingly, an additional stage of expansion may be provided by
appropriate configuration of the carriers.

[0043] The carrier material may have a higher CTE than the CMC. With
increasing temperature, the carrier material will expand more than the
CMC. With increasing temperature, the greater thermal expansion of the
carrier will cause the gaps 170 to shrink (not merely in angular extent
but in linear dimensions).

[0044] Eventually, at a characteristic engine temperature T1,
associated with P1 in FIG. 6, the circumferential gaps will fully
close (an "initial lock-up" condition). Until this point, in a first
stage of the heating and speed increase, radial expansion has essentially
been due to the circumferential expansion of the CMC member 72 (e.g., the
member 72 alone has an effective weighting to account for at least 80% of
the expansion; as a practical matter, the example would be 100%). With
further increases in temperature, the carrier ring 90 will tend to expand
according to its CTE. This creates a second stage of thermal expansion
(see intermediate stage 554-2 of FIG. 6) wherein radial expansion is
dictated essentially or substantially by the carrier ring 90 and its CTE.

[0045] For example, if there is a relatively light radial compliance
between the carriers and the member 72, expansion of the carrier ring
against the compliant force (e.g., of springs 240) will progressively
close the radial play between the carriers and the member 72 and the
expansion will essentially be due to the carriers alone. With such
lighter compliance, the carrier ring 90 will essentially dictate further
thermal expansion during the stage 554-2 (with the member 72 playing no
significant role (e.g., substantially less than 20% and likely less than
5%). With much heavier compliance, the member 72 may go into noteworthy
circumferential tension placing the carrier ring 90 in corresponding
circumferential compression and somewhat countering its expansion. Thus,
such a heavy compliance stage may be substantially influenced by the
properties of both the carrier ring 90 and the member 72 (e.g., with each
of the two members having a weighted average contribution of at least
20%).

[0046] With further thermal expansion at an exemplary engine temperature
T2 associated with P2 the radial play is closed (entering the
stage 554-3 of FIG. 6) and further increase in temperature produces more
significant circumferential tension on the member 72 and compression on
the carrier ring 90. The properties of both may substantially dictate
further expansion subject to the specific properties and their relative
sizes. For example, if the longitudinal cross-section of the member 72 is
relatively large, its influence will be correspondingly large.

[0047] In certain possible configurations, T1 and T2 may be the
same so that the second interval is non-existent. In other possible
configurations, dimensions may be such that the play essentially never
closes and the third interval does not exist.

[0048] Additional considerations involve non-equilibrium operation of the
hybrid ring. These can be more complex than behavior of a single-material
ring. One area is hysteresis and differences in behavior on heating vs.
cooling. In the identified example, on heating, the high thermal
conductivity of the high CTE member causes it to heat up at a higher rate
than had it been made of a low CTE material. However, on cooldown, the
high CTE material will cool relatively faster. The low CTE ring thus
serves to slow shrinkage of the BOAS radius in decreasing power
situations, thereby protecting against pinch/rub in rapid reacceleration
situations. This is seen in FIG. 7 by the slight rightward offset of plot
564 from plot 560 after t4. Also, the reacceleration at t6 begins to
reheat the support ring. There is a brief interval after t6 (the initial
departure of 564' from 564) where the carriers have not yet locked up and
shallow expansion is due to expansion of the CMC. When the carriers lock
up, their CTE then dictates expansion along the steepened part of 564'
until the locked carrier ring either nears equilibrium or interferes with
the CMC.

[0049] The benefit is shown in FIG. 7, plot 564' whereupon a
re-acceleration the turbine rotor growth is less than the radial position
of the BOAS. Thus no detrimental contact between the blade tip and the
BOAS occurs. As the turbine rotor continues to grow as it gets hotter,
the control ring 72 has kept the carriers 92 at a larger radius, and
delays the initiation of the lockup. Once the carriers are hot enough,
lockup ensues, and the ring quickly moves back towards it's steady state
radius.

[0050] A noteworthy difference is the small difference in the radial
position of the BOAS during this transient phenomenon. The slow
responding ring 72 forces the carriers 92 to be held at a larger radius
during the transient deceleration and cool-down phase, even as they
un-lock. The small, extra residual outward radial displacement
constraint, typically 0.005-0.020 inch (0.13-0.5 mm) provide enough of a
difference between the hybrid assembly and the high CTE ring that the
transient pinch rub event does not occur, and no longer becomes the
limiter for turbine tip clearance.

[0051] Such a result may not likely be obtainable with a high CTE control
ring. Nor is it apparently obtainable on a bonded bi-material ring
(because the cool-down phase would cause the high thermal
conductivity/CTE inner portion to pull away from the low CTE outer
portion, thus the deceleration pinch point is similar to the high CTE
ring. Additionally, if one were to try to counteract this effect, large
radial tensile loads would have to be carried between the outer low CTE
ring and the high CTE inner ring, with significant structural challenge
of the restraint features and/or fasteners. In the hybrid design, the
segmented carriers 92 solve this problem because they can be held outward
by the ring 72 even as they unlock, and do not create any extra force on
the ring 72.

[0052] In yet an alternative embodiment, the carriers directly abut at the
initial speed. Effectively, the carriers (at least initially) behave as
if they were an integral structure. In such a situation, the initial
stage of thermal expansion is dictated by the CTE of the carrier while
the radial play between carrier and member 72 closes. There may be
sub-variations based upon the presence and properties of any springs or
other radial compliance. After closing of the play, the properties of
both the carrier and the member 72 are involved reflecting their relative
sizes. However, hysteresis described above opens up the possibility that
the carriers could separate on rapid cooling (e.g., quick deceleration)
thereby creating a temporary cooldown-only situation wherein only the CMC
member dictates BOAS radial position.

[0053] The carriers could be integrated (e.g., via carrier-to-carrier
fastening) or could be unitarily formed as a continuous full annulus
(e.g., wherein fore and aft carrier halves become full rings). There may
be sub-variations based upon the presence and properties of any springs
or other radial compliance. Such examples may further be divided into
situations where there is an initial radial gap vs. situations without
such a gap. Again, hysteresis opens up the possibility that a gap
develops only transiently on rapid cooldown, thereby creating a temporary
cooldown-only situation where only the high CTE material dictates BOAS
radial position.

[0054] One or more embodiments have been described. Nevertheless, it will
be understood that various modifications may be made. For example, when
implemented in the remanufacture of the baseline engine or the
reengineering of a baseline engine configuration, details of the baseline
configuration may influence details of any particular implementation.
Accordingly, other embodiments are within the scope of the following
claims.